System for producing remote sensing data from near earth orbit

ABSTRACT

A satellite system operates at altitudes between 180 km and 350 km relying on vehicles including an engine to counteract atmospheric drag to maintain near-constant orbit dynamics. The system operates at altitudes that are substantially lower than traditional satellites, reducing size, weight and cost of the vehicles and their constituent subsystems such as optical imagers, radars, and radio links. The system can include a large number of lower cost, mass, and altitude vehicles, enabling revisit times substantially shorter than previous satellite systems. The vehicles spend their orbit at low altitude, high atmospheric density conditions that have heretofore been virtually impossible to consider for stable orbits. Short revisit times at low altitudes enable near-real time imaging at high resolution and low cost. At such altitudes, the system has no impact on space junk issues of traditional LEO orbits, and is self-cleaning in that space junk or disabled craft will de-orbit.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation application of U.S. patentapplication Ser. No. 15/868,794, filed Jan. 11, 2018, entitled “SystemFor Producing Remote Sensing Data From Near Earth Orbit,” which claimspriority to U.S. Provisional Patent Application Ser. No. 62/430,727,filed Dec. 6, 2016, entitled “A Satellite System” and U.S. patentapplication Ser. No. 15/439,533 filed on Feb. 22, 2017, entitled “ASatellite System”. The entirety of U.S. Provisional Patent ApplicationSer. No. 62/430,727, U.S. patent application Ser. No. 15/439,533 andU.S. patent application Ser. No. 15/868,794 are incorporated herein byreference.

BACKGROUND

Satellites are used in many aspects of modern life, including earthobservation and reconnaissance, telecommunications, navigation (e.g.,global positioning systems, or “GPS”), environmental measurements andmonitoring and many other functions. A key advantage of satellites isthat they remain in orbit due to their high velocity that creates anoutward centripetal force equal to gravity's inward force. Therefore,once in orbit, they stay there typically for years or decades. Since thevelocities are so high (e.g., 3-8 km/s, depending on altitude),atmospheric drag should be minimized and/or avoided, which meanssatellites typically operate outside virtually any trace of theparticles that constitute the atmosphere. In addition to drag,atmospheric collisions with particles, even at trace concentrations, canoverheat, damage or eventually destroy the satellite.

Satellites are therefore clearly differentiated from atmospheric flying(i.e., airborne) vehicles such as airplanes, unmanned aerial vehicles(UAVs), helicopters or balloons, in which the atmosphere supports liftand the vehicles operate at velocities typically between zero (i.e.,hovering) to 1-3 times the speed of sound and at altitudes below about35 km.

Satellite orbital heights are typically categorized in three broadsegments: low earth orbit (LEO), medium earth orbit (MEO) andgeostationary earth orbit (GEO). The general uses and characteristics ofthese orbits are shown in Table I and represent generally accepted usageof the terms LEO, MEO and GEO. Satellites can orbit at any altitudeabove the atmosphere, and the gaps in altitude shown in Table 1, such asbetween LEO and MEO, are also used, if less regularly. It is also commonthat satellites may orbit in eccentric, non-circular orbits, therebypassing through a range of altitudes in a given orbit.

TABLE I Typical characteristics of common orbits. Altitude, Velocity,Orbit km km/s Exemplary Uses Comments LEO   400-2,000 6.9-7.8 Earthobservation, Random orbits, 3-10 Y sensing, ISS, telecom lifetime, spacejunk issue, constellations little radiation MEO 15,000-20,000 3.5 GPS,GLONASS, Highest radiation (Van Earth observation Allen Belt),equatorial to polar orbits GEO 42,000 3.1 Sat TV, high BW Can remainabove same telecom, weather spot on Earth, typically satellitesequatorial orbits

For most satellites, their useful lifetime is determined by multiplefactors. For example, in the case of GEO satellites, small fluctuationsin solar winds and earth's gravity require regular use of fuel tomaintain the satellite's position and attitude. Once exhausted of fuel,a satellite is typically rendered useless and decommissioned. However,due to GEO height, such a satellite itself will stay in orbit virtuallyforever due to its altitude and near zero atmospheric drag. Due to theirapparent stationary position as viewed from earth's surface, they arewidely used for telecommunications and satellite TV. Their largedistance from Earth limits their usefulness in telephone services (timedelay) and in high-resolution imaging (distance). They encounter solarwinds and cosmic radiation that force use of very specialized andexpensive electronics to survive.

MEO satellites are in the mid-range, mostly similar to GEO satellitesexcept that they do not appear stationary when viewed from earth'ssurface. Their most common usage is for satellite positioning services,such as GPS, and certain Earth observation missions for which theirtrade-off in altitude between GEO and LEO is beneficial. Due to thepresence of the so-called Van Allen Belts, these satellites can sufferlarge amounts of radiation and therefore require very specialized andexpensive electronics to survive.

LEO satellites, conversely, may be in a constant state of very slightatmospheric drag requiring either regular boost to their altitude (e.g.fuel burns of typically chemical engines) or an end-of-useful-lifecaused by reentry and burn up similar to a meteor entering the earth'satmosphere. As an example, the International Space Station (ISS),orbiting at about 425 km, loses approximately 2-4 km/month of altitudeand requires regular fuel burns to ensure it stays in proper orbit. Butthe atmospheric drag is still very low and LEO satellites can remain inorbit for years without fuel burns.

This relatively long life is the source of so-called “space junk”, inwhich any orbiting device can potentially collide with a usefulsatellite, thereby damaging or destroying it and creating additionalorbiting objects. It is a widely recognized issue that at some densityof space junk, probabilities of collisions increase, eventually leadingto a virtually unusable orbit. A beneficial element of the currentinvention is to provide satellite services without increasing the spacejunk issue and furthermore to enable a mechanism that will be“self-cleaning” in the chosen orbits of 180-350 km.

Due to various shielding effects, especially of earth's magnetic fields,LEO satellites encounter little radiation and therefore do notnecessarily require specialized and expensive electronics to survive. Anexception to this rule is the so-called South Atlantic Anomaly, or SAA,which is a region in which a higher density of energetic particles maybe found, causing short term interruptions of some electronics. Thiseffect can be mitigated by many known techniques, so does not present alarge issue for LEO satellites.

In fact, continual improvement in system operation is realized since bylowering the operating altitude, system components (e.g. optics,electronics, synthetic aperture radar (SAR), required solar panel area,etc.) can be made smaller, which in turn reduces vehicle size and drag,thereby enabling an even lower operating altitude, and so-on. While itis desirable to be closer to earth's surface (or any celestial body'ssurface, say Mars), atmospheric density effectively sets a lower limiton orbital altitude; or forces expensive, heavy counteracting systemssuch as on the Gravity field and steady-state Ocean Circulation Explorersatellite (GOCE), discussed below. For bodies without an atmosphere,such as earth's moon, there is no lower limit other than hitting thebody itself.

SUMMARY

The present disclosure relates, generally, to satellite systems in anear earth orbit, and more particularly to a satellite system capable ofhigh frequency, low latency data acquisition and transfer rates, thesystem occupying a near earth orbit.

As described below, a properly designed near earth orbit (NEO) vehiclemust generate thrust to overcome the vehicle's drag on a regular basis.As used herein, Near Earth Orbiters (NEOs) describe the system and itsconstituent vehicles (i.e., a “NEO satellite system”, “NEO vehicle” or a“NEO satellite”) that operate in stable orbits at 180-350 km (e.g.,below a typical LEO). Therefore, it is a purpose of this invention todescribe a satellite system based on orbital vehicles operating instable Earth orbits at altitudes well below traditional satellites,specifically between approximately 180 and 350 km.

The satellite system described herein employs a plurality of spacecraftconfigured to communicate with each other, as well as with terrestrialbased receivers (e.g., ground and/or sea based antenna). In someexamples, each spacecraft balances a variety of systems for sustainedoperation in a near earth orbit. For instance, drag is directlyproportional to atmospheric density. Therefore, each spacecraft has arelatively small cross-sectional area facing the direction of travelcompared to the total surface area, to reduce drag from collisions withatmospheric particles (e.g., oxygen, nitrogen, etc.). Each spacecrafthas a volume sufficient to support data collection equipment (e.g.,imaging and/or radar apertures), and a total surface area that is largerelative to the direction of travel surface area support solar energycollection. Each spacecraft contains a means of propulsion, which caninclude an engine and/or a volume of engine propellant (e.g., compressedxenon housed in tanks).

Further, each spacecraft includes one or more surfaces with solar energycollection panels, to provide power to a rechargeable battery and/or topower one or more components of the spacecraft directly. In someexamples, the solar panels are arranged about the spacecraft such thatsunlight is collected from various angles while maintaining a staticposition of each solar panel relative to the spacecraft bus. In thismanner, the solar panels act as passive aerodynamic control (i.e.,“stability fins,” in a swept-configuration, described in detail, below,with respect to one or more of the figures).

To achieve the multiple goals with a single vehicle, an examplespacecraft is defined by a thin, long bus, with stability fins extendingfrom the bus such that solar panels incorporated thereon are exposed tosolar energy regardless of the angle of the spacecraft. In someexamples, a sharp leading edge with a specialized coating may also beincorporated onto the spacecraft to reduce atmospheric drag. Further,the bus is capable of housing multiple components, including, but notlimited to, transceivers, processors, imaging systems, positioningsystems, and propulsion systems. In some examples, each spacecraft iscapable of maintaining an orbit of about 220 km to 280 km from Earth orgreater, for approximately three (3) years.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows an example satellite in accordance with aspects of thisdisclosure.

FIG. 2 shows a perspective view of an example satellite in accordancewith aspects of this disclosure.

FIGS. 3A-3C show example satellite configurations in accordance withaspects of this disclosure.

FIG. 4 shows an example satellite deploying a solar collection system inaccordance with aspects of this disclosure.

FIG. 5 shows another example satellite in accordance with aspects ofthis disclosure.

FIG. 6 shows a cross-section of an example forward surface of asatellite in accordance with aspects of this disclosure.

FIG. 7 shows a cross-section of another example forward surface of asatellite in accordance with aspects of this disclosure.

FIG. 8 illustrates graphical data of air particle interaction within anexample satellite in accordance with aspects of this disclosure.

FIG. 9 shows a cross-section of an example satellite illustratingvarious components in accordance with aspects of this disclosure.

FIG. 10 shows another cross-section of an example satellite illustratingvarious components in accordance with aspects of this disclosure.

FIG. 11 shows another perspective view of an example satellite inaccordance with aspects of this disclosure.

FIG. 12 shows an example of satellites interacting with plural groundstations in accordance with aspects of this disclosure.

FIG. 13 shows an example of satellite necklaces in accordance withaspects of this disclosure.

FIG. 14 provides a flow chart of an example method 260 of generating animage from data acquired by an imaging system of one or more NEOvehicles 100, in accordance with aspects of this disclosure.

The several figures provided here describe examples in accordance withaspects of this disclosure. The figures are representative of examples,and are not exhaustive of the possible embodiments or full extent of thecapabilities of the concepts described herein. Where practicable and toenhance clarity, reference numerals are used in the several figures torepresent the same features.

DETAILED DESCRIPTION

This detailed embodiment is exemplary and not intended to restrict theinvention to the details of the description. A person of ordinary skillwill recognize that exemplary numerical values, shapes, altitudes,applications of any parameter or feature are used for the sole purposeof describing the invention and are not intended to be, nor should theybe interpreted to be, limiting or restrictive.

The current disclosure relates to vehicles operating at altitudesbetween about 180 km to 350 km, what is defined herein as a Near EarthOrbiter (NEO), using thrusters for orbiting where atmospheric density istoo high for traditional satellites and too low for airborne vehicles.To remain in stable orbit, the NEO satellite employs a propulsion system(e.g., electric or chemical propulsion engine) to generate thrustsufficient to overcome the vehicle's drag. An exemplary NEO satellitebased on electric propulsion is described herein, but a person ofordinary skill in the relevant art will appreciate that other forms ofpropulsion are possible, such as chemical, optical or others that may besubstituted and align with the scope of the present invention. The NEOsatellite may be powered by solar energy and equipped with various datacollection systems, data transmission and reception systems, datastorage, power storage, as well as other systems.

The NEO satellite may include an attitude and orbit control subsystemconsisting of sensors to measure vehicle orientation, control lawsembedded in the flight software, and one or more actuators (e.g.,reaction wheels, thrusters, etc.). These actuators apply the torques andforces needed to re-orient the vehicle to a desired attitude, keep thesatellite in the correct orbital position, and keep antennas and sensorspointed in the right directions.

Additionally or alternatively, a thermal control subsystem can makeadjustments to protect electronic equipment from extreme temperaturesdue to intense sunlight or the lack of sun exposure on different sidesof the satellite's body (e.g. optical solar reflector). Another systemis the communication payload, which is made up of transponders. Atransponder is capable of receiving uplinked radio signals from earthsatellite transmission stations (e.g., antennas; see FIG. 12),amplifying received radio signals, sorting the input signals, anddirecting the output signals through input/output signal multiplexers tothe proper downlink antennas for retransmission to earth satellitereceiving stations (e.g., antennas).

Some example satellites disclosed with respect to the current inventionmay orbit at altitudes below LEO (i.e., lower than 350 km). Due toatmospheric drag at these altitudes, thrust may be provided on acontinuous or regular periodic basis or the vehicle's orbit will decayin a matter of days, weeks or months, depending on altitude. The NEOvehicle 100 described herein could be configured to provide sufficientthrust to maintain orbits between 180-350 km.

At altitudes lower than LEO, atmospheric density increasesexponentially. Below an altitude of about 120 km, atmospheric densitythat causes atmospheric drag increases by an order of magnitude aboutevery 20 km. Meanwhile, above that breakpoint and up to about 400 km,the atmosphere changes by an order of magnitude about every 50-75 km.The key effect is that atmospheric density, and therefore drag, is aboutfive orders of magnitude higher at an altitude of 100 km compared to thealtitude of the ISS at about 425 km. Furthermore, the high velocitycollisions with residual atmospheric oxygen are highly destructive tovehicle components at this range. Accordingly, very few satellitesoperate below about 400 km, and those that do are often in highlyelliptical orbits, thus spending very little time at the loweraltitudes. Conventional satellite design assumes that the vehicle designitself has little or no effect on drag, which is a fair assumption forsatellites operating at ISS altitudes and higher. Because of the rapidincrease in drag below these altitudes, it has conventionally beenthought that orbits in the zone of rapidly increasing drag had littleutility, and systems have conventionally been designed to avoid theselower altitudes. Certainly, traditional satellite design is poorlysuited to a high drag environment.

However, maintaining a lower altitude orbit is desirable for multiplereasons. For instance, any earth imaging application can get higherresolution images from a smaller, less complex imaging device simply bybeing closer to the surface. For example, if an optical imager is 3times closer to its object, it will get approximately 9 times betterresolution (i.e., in pixels per area) for a given optical system.Similarly, for telecommunications, due to the inverse square lawrelationship between radio frequency (RF) energy and distance, atransmitter that is 3 times closer will create 9 times stronger signalat a receiver, or require 9 times less power to achieve the same signalpower at the receiver. Additionally, for an active radar application,being 3 times closer requires 81 times less power for equivalentperformance due to the 1/r⁴ power law of radar (or 27 times less powerfor SAR, due to the associated 1/r³ power law). All of these factorsenable the exemplary NEO vehicle to reduce the size and cost of a NEOsatellite system sufficiently to enable large satellite constellationsthat have short revisit times at affordable cost.

FIG. 1 illustrates, in a front view, an exemplary version of a NEOvehicle 100. The NEO vehicle 100 can further include an electricpropulsion engine 106 to generate thrust by, for example, ionizing andaccelerating a propellant gas to counteract drag, in order to maintainthe desired orbit. Additionally or alternatively, the engine 106 can becomprise an engine with chemical thruster capabilities, such as amicro-thruster type engine. Although represented as being external tothe vehicle bus 102, the engine 106 can be integrated within the bus102, shielded by one or more panels of the bus 100, and/or dimensionedto extend beyond a surface of the bus 102, in accordance with thepresent disclosure. One or more stabilization surfaces or panels 108 canbe employed, designed to enhance the stability of the NEO vehicle 100,as well as support solar paneling to collect power. The NEO vehicle 100is defined by a narrow cross section, as exemplified in vehicle bus 102.The bus 102 includes a first or top panel 110, a second or bottom panel112, and lateral sides 114 and 116. At the nose of the NEO vehicle 100is a leading edge 104, which is configured with a bevel to slope towardone or both the first or second panel 110, 112.

The example NEO vehicle 100 of FIG. 2 is shown in perspective view,illustrating a bevel 118 sloping from the leading edge 104 to the panel110. In some examples, the bevel 118 is angled at 20 degrees, andanother bevel opposite bevel 118 slopes toward panel 112. In someexamples, the angle is greater than or less than 20 degrees. Inexamples, the bevel 118 slopes at a first angle, whereas the oppositebevel slopes at a second angle different from the first angle.Furthermore, the bevel can slope at a constant angle on a flat surface,or can progress at a varying gradient toward the panels 110, 112. Othervariations on the surface of the bevel can also be employed, includingovoid-conical shape, pyramidal shape, etc., with the key feature beingthat the frontal area is sharply angle along the direction of travel. Asdescribed herein, the small cross section of the NEO vehicle 100, aswell as the sloping bevel from the leading edge 104, reduces drag on thevehicle 100 from atmospheric particles and aids in maintaining stableorientation in orbit.

FIGS. 1 and 2 illustrate a NEO vehicle 100 in a deployed configuration.FIGS. 3A-3C show example satellite configurations for storage and/ortransport, in accordance with aspects of this disclosure. For instance,FIG. 3A is a view from the top or bottom of the NEO vehicle 100, whereasFIG. 3B shows a perspective view from the rear of NEO vehicle 100. FIG.3C shows a stack of three NEO vehicles, which can be arranged in a tightgrouping for efficient transport on a launch satellite. Once the launchsatellite delivers the vehicle(s) to an appropriate orbit, theconfiguration changes to operational. In the example of FIG. 4, the NEOvehicle 100 deploys the support panels 108 (housing the solar collectionsystem). As shown, the panels 108 are secured or otherwise arrangedtightly against the NEO vehicle body 102 during storage/transport, suchthat a pair of panels 108 are unfolded from each of the first surface110 and the second surface 112. In some examples, a light baffle isfurther deployed once the vehicle is delivered in the proper orbit (see,e.g., FIG. 9).

Stabilization surfaces/panels 108 can be arranged to extend from thevehicle bus 102 to both minimize drag on the vehicle and to maximizeenergy capture by solar paneling mounted thereon. In the example of FIG.5, the NEO vehicle 100 includes four panels 108 arranged at an anglefrom the horizontal plane. This arrangement provides for maintainedexposure of at least two solar collection surfaces regardless of vehicle100 orientation relative to the sun 120. Thus, at a 45 degree angle, apanel 108 would have an effective radiation exposure in accordance withthe following equation:d×cos θ

where d is the length of the panel and θ the angle of incidence of thesolar rays relative to the panel normal direction. Although shown in aparticular arrangement, the panels 108 can be angled in any manner withrespect to the vehicle bus 102. Further, the panels 108 can be deployedin a fixed position, or can be movable to maximize exposure to the sun120. In some examples, movement of the panels 108 can be commanded by acontroller, such as to turn a motor, or in response to a sensorconfigured to track the intensity of the sun's rays. In some examples,the satellite bus panels 110, 112, 114, 116 can also support solarcells. Of note in this design is that in all cases, the large panels arepresented edge on to the direction of travel, which substantiallyreduces drag effects from the panels relative to presenting the solarsurface to the direction of travel, and indeed the small frontal crosssection of the vehicle shown in FIG. 1 relative to its overalldimensions and volume is an important aspect of the inventive design.

FIG. 6 shows a side-view of an example forward surface of a satellitewith a drag analysis represented as a vector. For instance, the leadingedge of object 130 experiences significant drag force, shown asparticles rebound at a vector V₂ in response to particles travellingtoward the object 130 at a velocity V₁. Object 132 experiences a lessdirect effect 134, as the direction of the reflected particles (shown asvectors 134, V₂) are directed at generally more diffuse angles thanobject 130.

Unlike traditional atmospheric drag cases for controlled flight in theatmosphere, drag at these altitudes is a function of particle density,speed, and the mechanism by which those particles interact with thesatellite vehicle. Some particles will “bounce” off the drag surface in“reflective” collisions, which do not result in the full transfer ofmomentum from the vehicle to the particle, and the particle retains alarge vector V₂ relative to its original pre-collision vector V₁. Othercollisions are “thermally accommodative”, meaning that the particle iseffectively trapped or absorbed by the surface, with a full transfer ofmomentum and energy that results in the particle leaving the surfacewith a small residual vector V₂ relative to the drag surface. The lattercollisions have a higher contribution to drag. Conventional orbitalvehicle design does not account for or design against the two types ofdrag inducing collisions.

The objects 130, 132 shown in FIG. 6 are representations of a typicalsatellite surface material that exhibits diffuse reflection and fullthermal accommodation drag proportional to (V₁-V₂), which isapproximately equal to V₁, as V₁ is much greater than V₂. FIG. 7, bycontrast, exhibits the interaction of a front edge/bevel of a NEOvehicle 100 with a surface treatment and/or selected material, asdescried herein. An example of such materials and possible applicationsare described in co-pending application No. 62/616,325 , entitled“Atomic Oxygen-Resistant, Low Drag Coatings And Materials,” to TimothyMinton and Thomas E. Schwartzentruber. For instance, the object 138 iscoated with an advanced material (e.g., such as SiOx), which results ina partial specular reflection (shown by vector V₂, 136) and low thermalaccommodation. The presentation of angled surfaces in the direction oftravel, on the one hand, plus the use of materials that exhibitreflective properties with respect to residual atmospheric particles, onthe other, result in drag on the object 138 (e.g., the leading edge andbevel of NEO vehicle 100) that is much lower, as V₁ can be comparable toV₂ in magnitude.

In one example, different surface treatments or materials are used forthe exterior of the satellite system. For example, atomically smoothmaterials such as sapphire or other polished materials may be used,where interactions with atmospheric particles would be similar tospecular reflection. Accordingly, the endurance of the NEO vehicle 100may exceed that of traditional satellites. In the case of a NEO vehicle100 operating with an ion engine, interaction with oxygen, and to somedegree nitrogen, may limit endurance of the system. Such issues may bemitigated significantly, for example, by proper choice of materials. Forexample, metallic elements, such as heavy, noble metals like gold do notoxidize and are less susceptible to sputtering than other materials. Newsynthetic materials or high strength ceramics may also be used.

In some examples, advanced materials are used to reduce drag fromatmospheric particles. For instance, the use of advanced materialscombined with a beveled (i.e. sharp) leading edge can reduce the dragexperienced by the satellite system by approximately a factor of two.Importantly, the mass and volume needed to maintain orbit of thesatellite system decreases, as much as half. Similarly, the missionlifetime for the system could be doubled. The improvement can also havean impact on safety factors for endurance and weight.

FIG. 8 illustrates graphical data of air particle interaction inaccordance with aspects of this disclosure. At altitudes of 180-350 km,earth's atmosphere is made up primarily of O, O₂, N and N₂. In anexample where the NEO vehicle 100 orbits the Earth at about 200 km, theNEO vehicle 100 has an orbital velocity of about 7.8 km/sec, these atomand molecule species have a velocity relative to the vehicle of the same7.8 km/sec. The drag analysis presented in FIG. 8 shows a molecular beamof oxygen atoms (O), traveling at approximately 8 km/s. The molecularbeam impacts the SiO2 surface at approximately a 20-degree angle. Thenarrow scattering distribution is centered at approximately 30 degrees.Accommodation is low, with an E_(out)/E_(in)=60-80%. Thus, theapplication of a SiOx coating and angled incidence has a meaningfulreduction in drag versus traditional materials, on the order of one-halfthe drag experienced by a diffuse material without a beveled (i.e.sharp) leading edge.

FIG. 9 shows a view of an example NEO vehicle 100 with the bottomsurface removed to expose various components therein. As shown in FIG.9, a radio frequency antenna 150 (e.g., a phased array) can be included.An example of such a system and possible applications are described inco-pending application Ser. No. 15/868,812, entitled “Radio FrequencyData Downlink For A High Revisit Rate, Low Earth Orbit SatelliteSystem,” to Daniel Nobbe and Ronald E. Reedy. A computing platform 152can include a processor, memory storage, and/or various sensor types.Attitude control gyroscopes and/or reaction wheels can be included. Abattery 154 or other storage system (e.g., capacitor, etc.) can be usedto store power collected by solar panels in order to, for example, powerthe various components and the electronic engine 106 of the NEO vehicle100.

One or more optical imaging systems/lenses 156,158 are also included(e.g., variable field of view, multispectral imaging, etc.). The lenses156, 158 are configured to have a thickness sufficient to providedetailed imaging (e.g., a 1 m resolution at NEO altitudes) yet thinenough to fit within the vehicle bus 102, along with the various othercomponents. A folded light path contributes to reduced thickness of anoptical assembly, while a radar assembly can be made from an arraysimilar to the radio phased array antenna. Additionally oralternatively, the imaging system can include a mechanical device tocontrol the orientation of the lenses 156, 158 to adjust the focus ofthe imaging system. A baffle 162 can be used to provide stability aswell as filtering stray light effects from non-imaged sources, supportedby one or more posts 164. Each spacecraft is configured with sufficientarea/volume to house one or more imaging systems, such as two cameralenses 156, 158, and one or more baffles 162. In some examples, a cameralens can be a 10 cm thick optical lens system, and a baffle external tothe vehicle bus is used.

Many aspects of the spacecraft have equal applicability for systemsconfigured for image capture (e.g., optical data collection) and radarcapable spacecraft. In some examples, considerations related to size ofthe vehicle, weight, drag, power demands, as well as propellant needs,may change based on these and other factors. For example, in someembodiments, the cross-sectional area for an imaging satellite isgreater than that for a radar capable satellite (e.g., about 5 cm thickvehicle bus for radar satellite, compared with about 10 cm thick tohouse the camera optics).

Additional and alternative components may be included in the NEO vehicle100, such as radar or radio components, sensors, electronics bays forelectronics and control circuitry, cooling, navigation, attitudecontrol, and other componentry, depending on the conditions of theorbiting environment (e.g., air particle density), the particularapplication of the satellite (e.g., optical imaging, thermal imaging,radar imaging, other types of remote earth sensor data collection,telecommunications transceiver, scientific research etc.), for instance.In some examples, the system can include one or more passive and/oractive systems to manage thermal changes, due to operation of thecomponents themselves, in response to environmental conditions, etc. Thecomputing platform 152 can be configured to adjust the duty cycle of oneor more components, transfer power storage and/or use from a given setof batteries to another, or another suitable measure designed to limitoverheating within the NEO vehicle 100.

A propellant storage tank 160 is coupled with the engine 106 to generatethrust to counter the forces on the NEO vehicle 100 from drag, or toposition the vehicle in the proper orbit. The present and desired orbitcan be compared and any adjustments can be implemented by the computingplatform 152. For example, based on sensor data, the computing platform152 can determine spatial information indicative of a current altitudeof the satellite, an orientation of the satellite relative to aterrestrial surface, and a position of the satellite relative to othersatellites. This data can be compared against a desired altitude,orientation or position. If the computing platform 152 determines anadjustment is needed, the electric propulsion engine 106 is controlledto generate thrust sufficient to achieve the desired altitude,orientation or position.

FIG. 10 shows another cross-section of an example satellite illustratingvarious components and representative dimensions for the NEO vehicle100, in accordance with aspects of this disclosure. For instance, thevehicle, from leading edge 104 to the far end of the vehicle bus 102, isshown in the example of FIG. 10 as being approximately 120 cm long.Further, from the bottom edge of the engine 106 to the baffle 162 isapproximately 20 cm. As shown, the baffle 162 provides a filter foroptical imaging systems 156, 158. Moreover, a wide-angle reception bandof 45 degrees is illustrated for RF antenna 150. Additionally, FIG. 10shows a profile of the leading edge 104 and a top bevel 118 and a lowerbevel 119.

FIG. 11 shows another perspective view of an example satellite. Asshown, the baffle 162 is placed between the vehicle 100 and the surfaceto be imaged (e.g., toward Earth). In this example, two openings areprovided to accommodate two imaging systems. However, a single imagingsystem, or three or more imaging systems, are considered for thedisclosed vehicle. Although a one to one correspondence in imagingsystem to baffle opening is shown, a single baffle opening may be usedfor multiple imaging systems, or no baffle may be used, depending on theparticular application. Although not shown in FIG. 11, the baffle may beattached to the satellite bus in a number of mechanical ways, and thedeployment of the baffle may be performed in a number of mechanicalways, for example with a system of springs and latches. Additionally,the center of mass of the satellite should generally be located forwardof the center of pressure (e.g., the location where the net aerodynamicforce, due to particles impacting the satellite surfaces, acts).Locating the center of mass forward of the center of pressurefacilitates passive stability and aids in avoidance of any tumblingmotion. Similar considerations are engineered into design of passengeraircraft. The precise location of the center of mass and pressure can bearranged in many ways based on engineering tradeoffs.

In some example imaging systems, a baffle can be used to block a portionof incoming light. Accordingly, only light associated with the imagedsurface is transmitted to a lens of an imaging system. The baffle systemcooperates with the thin lens and imaging system to provide transmittedlight from the imaged surface to a detector (e.g., an array,photodetector, etc.) to collect data and/or images associated with theimaged surface. In accordance with the NEO vehicle described herein, thelens and imaging system are of a thickness sufficient to be fully housedwithin the vehicle bus (i.e. thinner than the bus height).

Moreover, the satellite systems described herein can employ variousforms of electric propulsion devices (e.g., ion engines) such as pulsedplasma thruster (PPT), Hall-effect thruster (HET), microwave discharge,and RF discharge devices. For example, Xe or Ar, both noble gases withrelatively high atomic masses, can be used as a propellant. Noble gasesare selected because they tend not to damage engine components, andmassive atoms efficiently convert energy into momentum. An exampleelectric propulsion engine is manufactured by PhaseFour, Inc.Additionally or alternatively, the engine 106 can comprise an enginewith chemical thruster capabilities, such as a micro-thruster typeengine.

FIG. 12 shows an example of satellites interacting with plural groundstations in accordance with aspects of this disclosure. As shown, aplurality of satellites 100A-100C are in a near earth orbit, asdescribed herein. A vehicle-to-vehicle laser communication system may beincluded to improve data download rates, flexibility and reliability.Each satellite 100A-100C is equipped with communications systems tocommunicate with other satellites (e.g., laser communications, radiocommunications, etc.).

For example, satellite 1008 can send and receive information tosatellite 100A via link 244 and with satellite 100C via link 246. In ahigh volume constellation with close spacing at low altitudes, line ofsight laser communications to neighbor vehicles will be effective. Inthe example of 90 satellites in an orbital plane at 1-minute intervals,distance between satellites will be approximately 450 km. Since thehorizon from 180 km altitude is more than 1,000 km away, a lasercommunications system is capable of providing a direct link to multiplesatellites in the same orbital plane with minimal atmospheric diffusioneffects at low power. Since the vehicles will be oriented along theorbital plane in order to minimize drag and their relative positionschange very slowly, the pointing system for the inter-vehicle lasercommunications may be relatively simple. Using such an inter-vehiclelink would enable very high-speed data rate transfer between vehicles,enabling downloads to be handled by a vehicle other than the onecollecting an image. Adding this flexibility to the system has severalbenefits, including filling dead-zone gaps, backup capability ifreceivers are unavailable, and backup capability if a downlinktransmitter on a NEO vehicle becomes disabled.

Although three satellites are shown in succession, any number ofsatellites into the tens of thousands can be employed in a satelliteconstellation, and can be aligned in a single direction of travel in asingle orbit, or may be traveling at angles with respect to each other,and occupy multiple orbits (see, e.g., FIG. 13).

As shown, each satellite 100A-100C is configured to send and receiveinformation to and from ground based systems 240A-240C. Each groundbased system 240A-240C is configured to communicate with another groundbased system via communication links 248, 250. For example,communications links 248 and 250 can be laser based, radio frequencytransmissions, wired or fiber optic connections, or a combinationthereof. The communication links may utilize dynamic beam shapes tomaximize data download during each pass of satellites.

The system further includes a distributed earth receiver system relyingon a large number of receivers each downloading data during a satelliteoverpass. For instance, ground based systems 240A-240C are configured tocommunicate with satellites 100A-100C to send and receive informationvia communication links 242A-242C. Additionally or alternatively, aground based system can communicate with more than one satellite, orvice versa. As shown in FIG. 12, ground based system 240A iscommunicating with satellite 100A via communications link 242A, and isalso configured to communicate with satellite 1008 via link 252. Inexamples, ground based station 240A can anticipate the arrival ofsatellite 1008 and adjust one or more antennas to facilitate datatransfer. The position of satellites within the orbit can be determinedbased on information stored in a database and available to each groundstation and/or satellite. The database can be updated in response todata received through earlier ground based station communications toimprove estimates of a given satellite's location, speed and/or otheroperational parameters. Moreover, communication between the ground basedstation 240A and satellites 100A and 1008 can occur simultaneously or insuccession.

It is a key element of the current invention that phased array antennasmay be used for radio communications, with antennas located on eachsatellite as well as on each ground based system. For example, phasedarray antennas permit variable antenna beam shapes to facilitate bothsignal acquisition (e.g., larger bean width with lower data rates) andsignal transmission (e.g., narrower beam width with higher data rates),as described in co-pending application entitled “Radio Frequency DataDownlink For A High Revisit Rate, Low Earth Orbit Satellite System”.

The example NEO satellite system described herein is capable ofproviding imaging, communication services, earth measurements, and othersatellite services based on one or more NEO orbiting vehicles operatingin long term, stable orbit at altitudes between approximately 180-350km. Further, the satellite system includes an array of such NEOsatellites in sufficient density to enable near-real time coverage ofthe earth. Benefits of the NEO vehicle 100 with a sustainable orbitwould accrue to virtually all other satellite applications, such ascommunications.

Short revisit times can be described as “near-real time.” TraditionalLEO and MEO satellites have revisit times from hours to days to weeks,depending on the number of satellites in the constellation. Due toextremely high satellite costs plus high launch costs, satelliteconstellations are typically limited to a few to a few dozen satellites.Some proposed systems include up to about 100 satellites, promisingrevisit times down to a day or so.

Near real-time revisit rates with a massive constellation offers manyadvantages and solves many problems inherent in current satellitesystems. One example is the “worst case” revisit time as compared to theaverage revisit time. Most satellites spend about half their orbit inearth's shadow (i.e., night) resulting in poor or useless images. Addingin cloud cover, up to 70% of earth's surface, sand storms andperspective issues (e.g., images taken around noon cast no shadow andare therefore more difficult to interpret) reduce the number of usefulimages to about one fifth or less of all images taken.

This sampling problem makes it difficult to plan image capture of acertain spot at a certain time. For many implementations, the averagetime to a useable image may not be as important as the worst case time,which we define as the time between images that meet a certain set ofcharacteristics (e.g., at a specific location, with specific lightingconditions, with specific weather, during a specific window associatedwith a specific event, etc.) In this example, getting images of aspecific area (e.g., a battlefield or a river flood plain) with a longrevisit time constellation can make a worst case scenario push from daysinto weeks. In this example, a system with a 3-day average revisit timecould be overhead at night for several sequential passes, and thenencounter cloud cover or dust storms when it is finally overhead withcorrect lighting. Therefore, an average revisit time of 3 days canbecome a one or two-week worst case scenario, a delay that reduces oreven eliminates the value of the images.

Conversely, with an exemplary revisit time of minutes (e.g., less thanan hour, to 10 or fewer minutes), the current NEO system will generallyhave a vehicle overhead any spot on earth during daylight hours, manytimes every hour. Furthermore, as clouds and dust storms are notstationary, the probability of having a NEO vehicle 100 overhead duringa break in the weather is further increased. Since these statistics arenot a purely linear extrapolation of the average revisit times (i.e.,they are exponential), worst-case revisit times become much moremanageable with the described low revisit time NEO system.

Images are only useful once they are conveyed back to systems on Earth.The NEO vehicle 100 includes a widespread array of receiving stationsrather than the normally low number of centralized receiving stationsfound in use with traditional satellite systems. For example, with threereceiving stations (e.g., US, Australia and Europe), a traditional LEOsatellite will be within transmission range approximately every 30minutes (90/3), at best. If the imagery data is available with aninherent delay of a week due to the long revisit time described above, afurther 30-minute delay is relatively small.

However, for the current NEO satellite system, with average revisitrates of less than an hour and down to minutes, such a delay would be alarge percentage of and possibly greater than the goal. Therefore, datacan be downloaded from the NEO vehicles to a large network of low-costearth receiving stations to enable low-latency data downloads, ideallywith latency from time of taking to time of receiving on the order ofminutes to tens of minutes.

In one exemplary solution, receiving stations may be mounted atopcommercial cellular base stations, of which there are about 300,000 inthe US alone. Most such base stations are designed to support cellularcommunications radially outward, not upward. Therefore, an upwardlypointed radiation pattern can use the open area at the top of the basestation tower, directing and receiving all energy to/from an orbitingNEO satellite and away from any interference with the cellular signals.

In order to download sufficient data during an overpass of a single NEOsatellite and to meet the size, mass and cost targets of the NEOsatellite, a simple antenna with a relatively wide beam will enable arelatively large footprint on earth's surface. For example, a beam withfull width half max (FWHM) beam angle of 45° from 100 km altitude wouldhave a circular footprint about 200 km in diameter. Assuming thevehicle's orbital velocity is about 7.8 km/s, a useable receive time ofabout 60 seconds would result. A narrower beam would reduce this timewhile a wider beam would increase it.

A tradeoff in the beam width is that as the beam width is reduced, themaximum data rate would typically increase. Hence, a tradeoff is made tooptimize how much data can be downloaded during a single pass over agiven receiver. A further improvement can be made by using a more highlyfocused beam on the receiving site and having it track the NEO satelliteas it passes overhead. This would enable a relatively long dwell timedue to the relatively broad NEO satellite transmission beam, along withrelatively high data rates due to the relatively tight receiver beam.Also, since mass of the receiver is not as critical as mass on the NEOsatellite, placing a more complex (i.e., heavier) receiver and trackingantenna on the receiving side will reduce overall system cost. In anexemplary embodiment, a NEO vehicle may collect images within a zone inwhich it is downloading those images to one of the plurality of groundbased stations. In this embodiment, images will download to acorresponding ground based station in virtual real-time (e.g., with adelay measured in seconds). As the NEO vehicle exits the zone ofacceptance for a first ground based station, it will hand off to anotherground based as the NEO vehicle travels, in a manner similar to cellularphone handoffs between cellular base stations

In order to ensure low-latency downloads, downloads may occur when avehicle is passing over long stretches of ocean or other “dead zones”,of which the oceans are the largest. In addition to ensuringavailability of sufficient receiving stations on islands, receivers mayalso be placed on ships, buoys, or platforms to receive the images,which can then be transmitted to processing centers via traditional highcapacity data links.

The described NEO vehicle 100 that maintains a stable orbit between180-350 km can be part of an array of satellites in an orbital plane,defined as a satellite necklace (e.g., a single orbital plane withmultiple NEOs). Ninety NEOs in a single polar necklace will enable oneof these satellites to traverse a given line of latitude about once perminute in a northbound direction assuming orbital times of about 90minutes, and again on the opposite side of the earth in a southbounddirection. In an example, twelve such satellite necklaces may bearrayed, each separated by one hour of longitude, may be able to imageany spot on earth on average about once per hour.

In the example of FIG. 13, one or more NEO vehicles 100 can maintain anorbit 256 around the Earth 258, in accordance with the presentdisclosure. In one example, 90 satellites per necklace can be used,however more or fewer satellites per necklace may be appropriate for agiven application. For example, 45 satellites per necklace would spacethe vehicles at 2-minute intervals, while 180 would space vehicles at30-second intervals. As a person of ordinary skill will understand, theearth will rotate during the interval between arrivals of two sequentialNEOs, with that distance determined by the time separation between thesatellites. Different spacing distances may impact other subsystemdesigns such as optical imaging and radio links, but the concept remainsthat a NEO satellite system can provide relatively high rates ofcoverage.

Since the time to revisit the same spot on earth is determined by thetime for that spot to rotate under the next necklace, doubling thenumber of satellite necklaces would reduce the revisit time for any spoton earth to 30 minutes or less, depending on the field of view of theonboard imager or radio. Conversely, halving the number of necklaces to6 would double revisit times. And reducing the number of necklaces to 4would triple revisit times. These changes would reduce system cost andcomplexity, which may be a reasonable tradeoff for certain applications.

Other orbital planes can be utilized, and non-uniform distributions ofNEO vehicles could have beneficial applications. For example, non-polarorbits would increase the amount of time spent over populated areas andreduce the amount of total time spent over the poles. In the initiallydescribed system of 12 satellite necklaces with 90 NEO vehicles pernecklace, twelve NEO vehicles would fly over each pole every minute.Arrays of necklaces with inclinations to the equator of less than 90degrees could provide shorter revisit times for areas of greaterinterest. A system of 48 near-polar necklaces with six vehicles pernecklace would yield a 15-minute global revisit rate, and higherdensities are contemplated for the presently described systems.

Additionally, sequences of satellites with shorter distances betweenthem in a given necklace may be better suited for certain applications.For example, ten satellites separated by a few seconds could providesequential data on phenomena such as floods, fires or ice melting thatcould be useful in scientific understanding. A variety of NEO satellitedistributions are possible for various applications, each of which canemploy the NEO vehicle 100 described herein. In some exampleconstellations, additional NEO vehicles in varied or relativelyrandomized orbits may reduce revisit times from an hour down to under 10minutes. In one example, ten thousand to twenty thousand NEO vehiclesmay provide coverage down to 1-3 minutes between images.

The NEO vehicle 100 may incorporate navigation, cooling, attitudecontrol, radio transmission, optical and radar imaging, power supplies,and digital processing. The resulting satellite can operate in longterm, stable orbits at altitudes between approximately 180-350 km, withthe capability to capture and transmit images of a given place on Earthon a high frequency basis, be it hourly or even more often.

Different altitudes, even variations of a few to more than tenkilometers, may offer different benefits for certain applications. Widerangle coverage from higher altitudes may be an adjunct to higherresolution coverage from lower altitudes. Combination systems in whichSAR radar is combined with optical images may be desirable to operatedifferent sensors (e.g. optical and radar) from different altitudes.

In another exemplary application, it may be beneficial to operate NEOsatellites operating at about 200 to 300 km to image orbiting space junkand satellites above them. Thousands of pieces of space junk, fromexpended launchers to small objects, represent a serious threat to LEOorbiting objects. Tracking such space junk from earth is difficult dueto their distance and atmospheric disturbance. Being much closer andmoving in independent orbits from those objects can improve trackingsubstantially. Thus, in some examples, data collection systems (e.g.,imaging and/or radar) can be oriented away from the Earth's surface. Inthis example, the satellite can house various collection systems thathave fixed and/or variable orientations, based on the desired scan.Accordingly, in order for a satellite in a NEO to collect data onsatellites in the LEO, the data collection system can be orientedopposite the Earth's surface.

Another advantage of the present invention is that the contemplatedorbits are “self-cleaning.” A NEO vehicle 100 deorbits from influence ofdrag within days or weeks without thrust, as will any debris from acollision. Deorbit times for vehicles still under control can beaccelerated by commanding the vehicle into a high drag orientation. Dueto the small size relative to conventional spacecraft, the NEO vehicle100 should entirely “burn-up” during reentry through the atmosphere. Asa result, no space junk is left in orbit when the satellite isdecommissioned, and there is substantially no risk of collisions withobjects and/or surfaces within the Earth's atmosphere.

For synthetic aperture radar (SAR) radar applications, a smaller numberof satellites per necklace and a smaller number of necklaces may besufficient to provide a desired frequency of useful images thanks to theall-weather and night imaging capabilities of SAR radars. SAR radarsemploy multiple transmission and reception antenna arrayed in a specificpattern. The pattern is typically longer in one dimension (e.g. thedirection of motion) than in the transverse direction. Therefore, arelatively rectangular array of elements may trail behind or be attachedto a NEO vehicle to provide the oblong radiation beam needed toconstruct SAR images. By trailing or attaching such an array of antennaelements, drag will be impacted only marginally since it will be in theparticle flow shadow of each NEO vehicle 100.

In the example NEO vehicle 100, due to the large savings in power, a NEOconstellation employing radar applications may create near-real timeradar imagery of the earth's surface. Considering a SAR as an example,typical satellite-based SAR systems in LEO orbits require averagetransmit powers in the kilowatt range. Such radars therefore requirevery large solar arrays to power them and then complex cooling systemsto remove the waste heat.

For a NEO SAR system with a 27 times reduction in power, the averagetransmit power consumption is reduced from, for example, 1 kW to about50 W. The solar panel size, weight and cooling required would alsoreduce by 81 times, thereby making such SAR systems that much cheaper tolaunch and operate. If the relative altitude were to be ¼ instead of ⅓of the traditional altitude, the savings would increase to 128 times andthe SAR example above may require less than 10 watts of transmittedpower.

The value and opportunity for this ultra-low power NEO SAR is that suchradars can image the earth's surface at night, through clouds, and eventhrough some dust storms. Therefore, a given NEO SAR system would beable to create useful images approximately 100% of the time whilereducing the statistical impact of night and cloud cover.

To achieve a SAR, an array of transmit/receive elements is provided withprecise spacing, typically at half the wavelength of the transmittedenergy. Such elements could be provided on a single NEO satellite withthe array attached to or trailing behind the NEO satellite in thedirection of motion, thereby creating the typically oblong beam patternrequired for SAR. The element array could also be created by a formationof NEO satellites that maintain accurate spacing, with such anarrangement also useful for longer wavelength radars. In both cases, therelative power savings is maintained due to the low altitude of the NEOorbits.

It is also possible to assemble arrays of NEO satellites positionedrelative to each other in a formation that may create the antenna arrayand beam pattern needed for SAR. This may be an optimal approach forlonger wavelength SAR applications since spacing between elements istypically related to the wavelength of the RF frequency being used.

The benefits and challenges associated with different orbits can beaddressed in response to a desired application's requirements.Combinations of SAR images with optical images provide uniquely usefulinformation. For example, radar may be able to image ground contoursthrough dense foliage that can be complemented by optical images of thefoliage. Different frequencies for both radar and optical imaging canalso add useful information. It is a benefit of the NEO system that nearsimultaneous imaging on a high revisit rate (e.g., hourly) providessubstantial improvements over traditional satellites at higheraltitudes.

In addition, a SAR-equipped satellite could be assigned to shadow anoptical imager, thereby providing tight correlation between radar andoptical images. Such a combination may provide a more comprehensiveunderstanding of activities on earth's surface than either type ofsatellite alone can offer. For example, a post-earthquake optical imagecan identify building damage that might be seen by terrestrial observerswhile radar images could highlight where vertical displacement hasoccurred in the building, or is occurring as a precursor to anaftershock. Such combinations today rely on long time lags between thetwo types of imagers, especially due to the scarcity of SAR-capablesatellites.

Orbital planes other than polar are possible as well as a hybrid mix ofpolar and non-polar planes. The specific orbital plane may be modifiedfor different applications. Land mapping satellites may be concentratedin lower latitudes since that is where the majority of earth's landmasses are found. The NEO vehicle 100 described herein can be applied toany orbit in the targeted altitude, from polar to equatorial.

Other Earth observation requirements also benefit from lower altitudeorbits. For example, the European Space Agency GOCE satellite configuredto provide highly accurate gravitational measurements, was placed in asomewhat lower orbit. The planned orbit was approximately 270 km, withoperational orbits achieved at 255 km and 235 km. To stay in the plannedorbit for the desired 3-year life, the satellite carried an ion thrusterto expel its stored Xe atoms, thereby creating sufficient thrust tocounteract the atmospheric drag. Launching sufficient Xe into orbit wasboth expensive and heavy. However, GOCE was much more massive(approximately 900 kg) than the contemplated vehicles of this invention(approximately 5-20 kg in one example), orbited at approximately 255 kmfor most of its design life, was not designed with the drag reducingfeatures of the invention, and was not designed to work as aconstellation providing low revisit rate earth sensing data. A majorearth observation opportunity for NEO vehicles such as disclosed hereincorresponds to information transfer (e.g., radio transmissions) for usein the Internet of Things (loT) technology. As electronic systems expandtheir direct links between unmanned systems (such as pipelines ordriverless vehicles) a NEO satellite may be very effective at monitoringsuch systems, independent of where the systems are located. Receivingradio transmissions and then retransmitting the information to the nextavailable receiver, such as a ground base station, may become animportant advantage in such autonomous systems.

FIG. 14 provides a flow chart of an example method 260 of generating animage from data acquired by an imaging system of one or more NEOvehicles 100, in accordance with aspects of this disclosure. In step262, an imaging system (e.g., imaging system 156, 158; radar 150; etc.)from one or more satellites (e.g., NEO vehicles 100) acquires data fromone or more sources (e.g. a surface of the Earth; another objectorbiting the Earth; etc.). In step 264, the data is transmitted to areceiver, such as at one or more ground stations. The receiver canreceive data from a single or multiple imaging systems on a NEO vehicle100. Additionally or alternatively, the receiver can receive multipleimages from multiple NEO vehicles 100, as described with reference toFIG. 12.

In step 266, the image data is processed. For example, the image datacan be compiled from multiple sources to enhance image quality, and/orto build an image of a larger area. Moreover, other processingassociated with signal transfer, such as buffering, can be performed atthis stage. Once data processing is complete, an image is generated inblock 268 based on the imaged data from the NEO satellite 100.

In summary, a satellite system has been described that comprises thefollowing features:

-   -   Satellite operation in stable orbits at altitudes from 180-350        km;    -   Vehicle arrays sufficiently dense to enable overflights on        approximately an hourly basis;    -   Solar powered electric propulsion engines on each vehicle        providing thrust to counteract drag;    -   Thin, low drag, resistant, materials and shapes that minimize        drag at orbital velocities and these altitudes;    -   High density array of receiving stations enabling low-latency        data downloads;    -   Laser-based vehicle-to-vehicle communications system;    -   Platform for optical and radar imagers;    -   Generate an image from one or more data capture systems;    -   Provide near real-time revisit rates from satellite to ground        based antenna; and    -   Self-cleaning orbital system for any generated space junk, to        name but a few advantages of the present disclosure.

A satellite system is described operating at altitudes between 180 and350 km relying on NEO vehicles. The system operates at altitudes thatare substantially lower than those in which traditional satellitesoperate, thereby reducing the size, weight and cost of the vehicles andtheir constituent subsystems, such as optical imagers and radio links.This reduction in size enables a virtuous cycle of further reduction invehicle drag, which enables lower altitude flight and further reductionin the size of vehicle components, etc., and makes massively largeconstellations of low cost, low mass vehicles feasible.

The system includes a large number of the low-cost, low-mass, lowaltitude NEO vehicles, thereby enabling revisit times substantiallyfaster than any previous satellite system. The NEO vehicles spendvirtually all of their orbit at the low altitude, high atmosphericdensity conditions that have heretofore been virtually impossible toconsider. Short revisit times at low altitudes enable near-real timeimaging at high resolution and low cost. The system further includes adistributed earth receiver system relying on a large number of receiverseach downloading data during a satellite overpass. The communicationlink may utilize optimized beam shapes to maximize data download duringeach pass. A vehicle-to-vehicle laser communication system may beincluded to improve data download rates, flexibility and reliability. Byoperating at such altitudes, orbital mechanics ensure no impact on thespace junk issues of traditional LEO orbits and the system isself-cleaning in that any space junk or disabled craft will quicklyde-orbit.

What is claimed is:
 1. A satellite configured to maintain an orbitbetween 180 km and 350 km, the satellite comprising: a vehicle buscomprising: a frontal section having at least one beveled surfaceconfigured to reduce drag; and a material applied to at least a portionof the at least one beveled surface to provide at least partial specularreflection of incident atmospheric particles, the material selected andthe at least one beveled surface configured to reduce drag on and damageto the material from interactions with atmospheric particles incident tothe satellite at during orbit by the partial specular reflection of theatmospheric particles.
 2. The satellite defined in claim 1, wherein theat least one beveled surface and the material are configured to reducedrag on the frontal section by more than 25 percent compared to the samefrontal section geometry with a material that results in fully diffusereflection.
 3. The satellite defined in claim 1, wherein the at leastone beveled surface and the material are configured to reduce drag onthe frontal section by more than 50 percent compared to the same frontalsection geometry with a material that results in fully diffusereflection.
 4. The satellite defined in claim 1, wherein the at leastone beveled surface comprises at least one of a) a conical shape, b) awedge shape, c) a pyramidal shape, or d) a curved shape.
 5. Thesatellite defined in claim 1, wherein the frontal section furthercomprises a leading edge that includes a first beveled surface of the atleast one beveled surface sloping toward a first panel of the vehiclebus and a second beveled surface of the at least one beveled surfacesloping toward a second panel of the vehicle bus opposite the firstsurface.
 6. The satellite defined in claim 1, wherein greater than halfof the frontal section comprises the at least one beveled surface. 7.The satellite defined in claim 1, wherein a mass of the satellite isless than 20 kilograms.
 8. The satellite defined in claim 1, wherein oneor more stability fins extend from the bus to provide aerodynamic flightstability.
 9. The satellite defined in claim 1 further comprising anengine configured to generate thrust by electric, mechanical or chemicalpropulsion sufficient to maintain the orbit between 180 km and 350 km.10. The satellite defined in claim 9, wherein the engine is configuredto generate propulsion by use of a Xenon propellant.
 11. The satellitedefined in claim 9, wherein the engine is a pulsed propulsion system.12. The satellite defined in claim 1, further comprising a solar energycollection system having a plurality of solar panels to provide power toat least one of the engine, a rechargeable battery, or a component ofthe satellite system.
 13. The satellite defined in claim 12, wherein thesolar energy collection system is electrically connected to theplurality of solar panels and configured to store energy collected bythe plurality of solar panels.
 14. The satellite defined in claim 12,wherein each solar panel is configured to deploy from a folded positionduring transport or storage to the outwardly extended position duringoperation.
 15. The satellite defined in claim 12, wherein each solarpanel is secured to the vehicle at about a 45-degree angle from one of afirst or a second panel of the vehicle bus in an outwardly extendedposition.
 16. The satellite defined in claim 12, wherein one or moresolar panels of the plurality of solar panels is oriented parallel tothe direction of flight.
 17. The satellite defined in claim 12, whereinthe at least one beveled surface is less than or equal to a 20-degreeangle relative to one or more surface panels of the vehicle bus.
 18. Thesatellite defined in claim 17, wherein each of the one or more surfacepanels of the vehicle bus has a width less than half of a length of thepanel, the vehicle bus further comprising lateral side panels betweenthe first and second panels with a height of the lateral side panelsless than half the length of the lateral side panels.
 19. The satellitedefined in claim 18, wherein when the satellite is configured forlaunch, the plurality of solar panels are folded onto the one or moresurface panels of the vehicle bus.
 20. The satellite defined in claim 1,further comprising sensors comprising one or more of an optical imager,earth sensor, a radio receiver, or a radar imager.
 21. An image or datafile generated from data acquired by at least one of the sensors andtransmitted from the satellite as defined in claim
 20. 22. An image ordata file received at the ground station, the image generated at thesatellite from data acquired by at least one of the sensors of thesatellite and transmitted from the satellite as defined in claim
 20. 23.A satellite configured to maintain an orbit having an altitude between180 km and 350 km, the satellite comprising a vehicle bus comprising: afrontal section having at least one beveled surface configured to reducedrag; a material applied to at least a portion of the at least onebeveled surface to provide at least partial specular reflection ofincident atmospheric particles, the material selected and the at leastone beveled surface configured to reduce drag on and damage to thematerial from interactions with atmospheric particles incident to thesatellite during orbit by the partial specular reflection of theatmospheric particles; an engine configured to generate thrust; and atransceiver configured to transmit or receive data from anothersatellite or at least one ground station.
 24. The satellite defined inclaim 23, wherein the satellite comprises a controller configured to:determine spatial information indicative of at least one of a currentaltitude of the satellite, an orientation or position of the satelliterelative to a terrestrial surface, and a position of the satelliterelative to at least one other satellite; compare the spatialinformation of the satellite against a desired altitude, orientation orposition; and control the engine to generate thrust sufficient toachieve the desired altitude, orientation or position.
 25. The satellitedefined in claim 23, wherein the satellite comprises an accelerometer, agyroscope, or a global positioning system.
 26. The satellite as definedin claim 23, wherein there are more than six orbital planes.
 27. Thesatellite as defined in claim 23, wherein each satellite is configuredto de-orbit within 90 days of loss of thrust.
 28. The satellite asdefined in claim 23 comprising 48 orbital planes each with sixsatellites per orbital plane.
 29. The satellite as defined in claim 23,wherein the satellite is configured to: transmit a first transmission toa first ground station of the at least one ground stations; and transmita second transmission to a second ground station of the at least oneground stations within 15-minutes of the first transmission.
 30. Thesatellite defined in claim 23, wherein the satellite comprises animaging system configured to collect optical images.
 31. The satellitedefined in claim 30, wherein the imaging system comprises one or morelenses with a thickness that is at least 80% a height of the vehiclebus, wherein the one or more lenses are oriented generally perpendicularto the direction of flight.
 32. The satellite defined in claim 30,wherein the one or more lenses are arranged to produce a folded lightpath.
 33. The satellite defined in claim 30, wherein the imaging systemcomprises a mechanical device configured to adjust an orientation of theone or more lenses.
 34. An image generated from data acquired by theoptical imaging system and transmitted from the satellite as defined inclaim
 30. 35. The satellite as defined in claim 23, wherein thesatellite occupies a first orbital plane, the satellite operating in anetwork of satellites, the network further comprising a second pluralityof satellites that occupy the first orbital plane or a second orbitalplane different from the first orbital plane.
 36. The satellite asdefined in claim 35, wherein the first and second orbital planescomprise at least 45 satellites in each plane.
 37. The satellite asdefined in claim 35, wherein the network comprises more than a thousandsatellites.